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RDTfcE PROJECT NO. 1F141807D174

USATECOM PROJECT NO. 4-6-0300-01 TESTING ACTIVITY PROJECT NO. 65-31

ENGINEERING FLIGHT EVALUATION OF UH-2B

ARMED HELICOPTER

FINAL REPORT

Mr. John N. Johnson Project Engineer

Mr. William A. Anderson Project Pilot

MARCH 1966 /

DISTRIBUTION STATEMENT A Approved foi public release]

Oittiibution Unlimited

U.S. ARMY AVIATION TEST ACTIVITY ■ DWANDS APB, CALIFORNIA

RETAIN IN CEMTRAL FILE SYSTEM UNTIL

DECEMBER 31, 1966, RETIRE JANUARY

1^67.

/

ENGINEERING FLIGHT EVALUATION OF UH-2B ARMED HELICOPTER

ROUE PROJECT NO. 1F1418070174

(USATECOM PROJECT NO. 4-6-0300-01) (USAAVNTA PROJECT NO. 65-31)

FINAL REPORT

JOHN N. JOHNSON WILLIAM A. ANDERSON PROJECT ENGINEER PROviECT PILOT

MARCH 1966

U. S. ARMY AVIATION TEST ACTIVITY

1ii

m

TABLE OF CONTENTS

ABSTRACT

SECTION 1. GENERAL

1. Background 2. Description of Materiel 3. Objectives 4. Summary of ResuVts 4J Cockpit 402 Engine Start 4.3 Taxi 4.4 Takeoff, Hover and Landing 4.5 Level Fllgh;. 4.6 Stabilized Turning Performance 4.7 Level Accelerations 4.8 Level Decelerations 4.9 Level Deceleration Side Flares 4.10 Turn Reversals 4.11 Tear Drop Turns 4.12 Autorotational Entries 4.13 Stability and Control 4 13 1 Static LonTUudinal SpceJ Stability 4 13 2 Static Lateral Directional Stability 4.13 3 Controllability 4.13 4 Maneuvering Stability 4J3 5 Dynamic Lateral Directional Stability

APPEf WICES

I. Test Data II. References

Paae

v

1

1 1 5 5 5 5 5 5 6 8 10 10 11 12 13 16 1ft 18 18 18 19 20

21 67

iv

ABSTRACT

This report presents the results of an engineering flight evaluation conducted to detemlne the technical engineering flight characteristics of the armed UH-2B helicopter. The evaluation was conducted by the U. S. Amy Aviation Test Activity (USAAVNTA), Edwards Air Force Base, CaMfomla. Tests were conducted at Edwards Air Force Base and Bakersfleld, California. Twenty-four productive flight hours were flown In the UH-2B, Bu No. 152202, during the period 17 November to 2 December 1965. The USAAVNTA was assigned responsibility for preparing the test plan, executing the teat and preparing the test report.

The maneuvering characteristics of the helicopter were stressed during the evaluation. Tests conducted included steady turns, pull-ups, accelerations and decelerations, tear-drop turns, and turn reversals. The maneuvering characteristics during an abrupt pull-up were undesirable. A negative maneuvering stability gradient occurred following a pull-up at high airspeeds and heavy gross weights'. This apparently occurred when the helicopter encountered blade stall. Directional control step inputs caused a nose-down pitch with right pedal and a small nose-up pitch with left pedal. This characteristic was undesirable for weapons firing.

■»Wb

SECTION 1 - GENERAL

1. BACKGROUND

The Departnent of th« Army directed the U. S. Army Materiel Command (USAHC) to conduct an expedited flight test evaluation of a selected group of armed helicopters. In messages, 23 October and 28 October 1965 (References a and b. Section 2, Appendix II), USAHC assigned this program to the U. S. Army Test and Evaluation Command (USATECOM) for testing by the U. S. Army Aviation Test Activity (USAAVN'A). under the technical direction of an AMC appointed representative who had full responsibility for the conduct of the flight test program. The Plan of Test of the Armed Helicopters (Reference c) was submitted by USAAVNTA 28 October 1965 and approved 8 November 1965. The test program was conducted at Edwards Air Force Base, California, from 17 November to 1 December 1965. An interim summary report of the combined armed helicopter test results (Reference d) was submitted by USAAVNTA on 6 December 1965 to the Chairman, Improved Armed Hel'opter Evaluation Group, Hq, USAMC. This report presents the final result f the engineering flight «valuation of the armed UH-2B helicopter.

2. DESCRIPTION OF MATERIEL

2.1 B.ASIC UH-2B HELICOPTER

The UH-2B helicopter, Bu No. 152202, was provided for this evaluation through the cooperation of the U. S. Navy, The helicopter has a single ma'<t -^tor and an antitorque tail rotor. The helicopter is powered by a T5B-GE-8 turboshaft engine. The engine is mounted above the cabin behind the main rotor mast. The four-bladed main rotor is driven through the main gear box which is mounted in front of the engine. Cyclic and collective pitch controls are obtained through blade servo flaps mounted on the main rotor blades. Aerodynamic action of the flaps stabilizes and changes the pitch of the main rotor blades, which are otherwise unrestrained in feather, in response to the pilot's operation of the cyclic and collective controls. An in-flight blade tracking system automatically adjuststhe tip path of the main rotor blades in response to 1-per-rev out-of-track vibration sensor signals. The aircraft is equipped with automatic stabilization equipment (ASE). This equipment, however, was not used during this evaluation. A yaw rate damper was Installed In the aircraft and was used for all stability and control and maneuverability tests. The aircraft has a retractable main landing gear and a full-swivel non- retractable tail wheel. The principal dimensions and general data are as follows:

Maximum Gross Weight 10,000 lb

Maximun Forward Center of Gravity (C.G.) " Sta. 158

Maximum Aft C.G. Sta. 173

Tt ' . 'if I a

Main Rotor

100X RPM 277 RPM

Disc Area 1520.5 sg ft

Total Blade Area (Including Servo Flaps) 134.2 sq ft

Blade Diameter 44.0 ft

Blade Chord 20.0 In

Solidity Ratio .0965

Shaft Angle-From Vertical Forward 6 deq

Tall Rotor

Disc Area 50.4 sq ft

Blade Area 7.0 sq ft

Blade Diameter 4.0 ft

202 UH-2B ARMAMENT

For the purpose of this evaluation the UH-2B was confiqured with a modified TAT 141 and two LAU 3A/A TO-round 2.7f.-1nch rocket pods. The TAT 141 consists of an M-5 greriade launcher and an XM-134 7.62 mini gun. The M-5 grenade launcher, however, was not installed for this evaluation; and Instead the test airspeed boom,which was considered to be equivalent in drag area to the launcher, was used. The modified IM 141 was mounted in the nose of the aircraft and the rocket pods were moupted under stub wings located on the side of the fuselage near the cabin area.

my Hiring u» nnif

FOR OFFICIAL USE ONLY

FOR OFFICIAL USE ONLY

3. OBJECTIVES

The objective of this test was to determine the technical engineering flight test characteristics of the UH-2B in an armed configuration.

A» SUMMARY OF RESULTS

4.1 cogpiT

The pilot station was comfortable. The seat and pedal adjustments were large enough to accommodate a variety of pilot dimensions. Both collective and cyclic controls were comfortable to operate throughout their range of travel. All other controls necessary to the mission were within easy reach. Visibility out of the cockpit was somewhat restricted for low- level maneuvering, particularly for steep left turns, during which the overhead console obscured a large area of the view. Vibration levels were generally low and caused no fatigue or distraction.

4.2 ENGINE START

The engine start procedure was quick and simple. Starts could be made witn intenial or external power. The engine was started with the rotor brake ON. With a fully charged battery the engine accelerated to qround-idle in less than 20 seconds, after which the rotor brake could be released. Subsequent engine and rotor acceleration characteristics were good. Ten seconds after release of the rotor brake, 10C-pcrcont rotor cpccd (277 RPM) could be reached. The yaw damper could be engaged approximately 2 minutes after rotor start. All checks could be carried out rapidly and short "scramble" times were possible. From crew entry to lift-off, a minimum elapsed time of 2 minutes could readily be achieved.

4.3 TAXI

Taxiinq was conducted using 90- to 100-oercent rotor RPf'. A small amount of collective pitch usually was reouired to initiate the taxi. The aircraft could easily be taxied rearward with the tall wheel locked. Durinq taxi forward small changes in direction (2 to 3 degrees) could be made with the tail wheel locked. The toe-ooerated wheel brakes were very effective. With the tail wheel unlocked, very tight ground maneuvering was possible and directional response to pedal was quick and effective. The overhead tail- wheel locking lever was sometimes awkward to operate. To release the locking pin required some pedal rocking unless the tail wheel was trailing symmetrically.

4.4 TAKEOFF, HOVER AND LANDING

Handling qualities and performance capabilities of the aircraft durinq hover, takeoff and landing were not quantitatively evaluated. Although there were no unusual characteristics, the following qualitative handling qualities were noted:

a. Vertical lift-off to a hover was normal and presented no unusual control problems. Hovering attitude was essentially level. Stability during hover was satisfactory for a single-rotor helicopter not employing stability augmentation. Aircraft response to cyclic» collective and pedal controls was satisfactory. A running takeoff was usually made as this was the quickest and easiest method of attaining cruise speed or dint speed.

b. The landing gear suspension was stiff and vertical landings were usually accompanied by some skipping and bouncing. Run-on landings were usually made. In these landings, the tail wheel was allowed to touch the ground during the flare before the main wheels were lowered.

4.5 LEVEL aiGHT

4.5.1 Level flight tests were conducted at density altitudes ranging from 800 feet to 6350 feet and gross weights ranging from 7600 pounds to 9690 pounds. Individual test results are presented In Section 2, Appendix I, Figures 4 through 11 and sumtarlzed In Figures 2 and 3. Limited range performance with the simulated TAT 141 and two LAU 3A/A 19-round rocket pods without nose cones Is presented In Figure 1. Rotor speeds of 277 (100 percent) and 266 (96 percent) were used for the tests conducted. There were no unusual level flight characteristics up to the maximum airspeeds tested. Level flight performance tests were also conducted to determine the effects of the addition of the rocket-pod nose cones and the removal of the left cabin door on power required, true airspeed and equivalent flat plate area. At a gross weight of 9300 pounds and 130 knots true airspeed (KTAS) reduced to sea-level standard- day conditions, a decrease In equivalent flat plate area of 1.2 square feet was realized with the addition of the rocket pod nose cones. At these conditions this was equivalent to an Increase of approximately 3 KTAS, or a reduction In power required of 25 shaft horsepower (SHP). Under the same conditions, with the removal of the left cabin door, an Increase of approximately 4.3 square feet In the flat plate area was realized. This was equivalent to a decrease of approximately 5 knots, or an Increase of 90 SHP required. The following graph shows engine output shaft horsepower versus true airspeed at 9300 pounds gross weight, and 100-percent rotor RPM, sea- level standard day:

:rp iT'iY * \ J * \ ^ ' * • V O L w t i L I

1300'

IC-AJ,

90 100 110 120 150 TRUE AIRSPEED - KNOTS

2 Ve r t i ca l and l a t e r a l v i b ra t i on cha rac te r i s t i c s were measured a t the j t ' s seat and in tiio cargo co.;part ;xnt d i r e c t l y below the t ransmission.

. . \ i ion data is presented in Figures 14 through 16. At gross weights of U ;ir,i. iiis to 9720 pounds, the 'JH-2U exh ib i ted low v i b ra t i on cha rac te r i s t i c s o . n r . g ]«";.•»i H i s i l t auovo 70 KCAS a t 100-percent ro to r speed. These c.. • s'.^cs wore w i t l i i n ti.e l i m i t s of Paragraph 3 .7 , MIL-H-8001A (;'o."»:rcnw e ) . Tiie">- ie r - rev v e r t i c a l s ing le-ar .p l i tudc v i b r a t i o n accelerat ions

ured in tiio cabin area, however, a id excoed MIL-H-8501A between 50 - 70 ..L.O i.liove '.'o/0 pounds gross weight . V ib ra t ion data was not recorded below

iro, . I r .atoly 50 KCAS. The a i r c r a f t ba t te ry was mounted as a 4-per- rev v.ui\.;W<n * i was tuned at approximately 100-percent r o t o r speed. ,\t .c . .w ro to r speeus, a marked increase 1n the 4-per-rev v i b r a t i o n was no t i ced . Tne automatic blade t rack ing system provided very good 1-per - rev ro to r t r i c k i n g i n level f l i g h t but was automat ica l ly disengaged dur ing

Fi),l CrriCiAL USE OiiLY

Hi , n i tiM

cr n r - j -?*: OMI V | V f i ^ I i x i i i h vVfc w l > L I

maneuvering f l i g h t above .15 g. Following any severe of several seconds was required for any resulting out-of-track condition to be corrected. One-g blade s ta l l was characterized by heavy 1-per-rev vibration and a s l ight buildup in 4-Por-rev vibration. The ^ " e r warn'Ing was reduced by the vibration absorber, which consequently reduced blade s ta l l warning. During maneuvers there was Inadequate olade s ta l l wacplng.

<i 5 3 Sound measurements were obtained at the copi 'o t 's head posi on using U* S Air Force Bureau of Standards calibrated equipment. The coc».,>it noise level was moderately high, dominated by a high-pitched transmission whine. I t should be noted, however, that a l l soundproofing had been removed. As shown in the following f igure, 1n level f l i gh t at • n

1a v e r ® 9 ® . f *

of 8130 pounds, recorded noise levels varied from a low of 111 ^ t e l s at 70 knots calibrated airspeed (KCAS) to a high of 118.5 decibels at 132 KCAS. The noise never readied uncomfortable or distracting proportions. Test data is presented 1n Figure 17.

zoo <0 150

5? 100 <0

& 50

0 40 60 60 100 120 CALIBRATED AIRSPEED ~ KNOTS

140

STA1.)iLIzrr> Tu;t:;inc PERFORMANCE

t, 6 1 Stabilized turning performance data was collected usir.g two methods. Tr.e"first method was to establish a trim speed at the test alt i tude and rocord the power required. Various stabil ized bank angles were estab ished while holding trim power and alt i tude constant and the resulting stabil ized airsnecds were recorded. Bank angle was increased unt i l alt i tude could no lor.rf.r uO maintained. The second method used was to perform the same mar.ouver maintaining the desired tr im sneed and alt i tude and adlustlng power nn«i bank angle as required. This maneuver was performed using various tr im speeds up to maximum airspeeds.

rul OFFICIAL USE ONLY

(;?ic;.VL USE ciM

4.6.2 Stabilized level turning performance tests using the first method proved to be Impractical In the UH-2B because of difficulty In maintaining trim power due to the fuel control governor Inputs as rotor loading varied. The second method was more suitable. The gener«! results and limiting factors obtained using this method are presented In the next paragraph.

4.6.3 In all cases, at the maximum bank angles, power available was the limiting factor, rather than rotor blade stall. Four trim speeds wore flown and cnch jpccd was maintained at Increased bank angles until full power was being used. Aft stick forces were high for a helicopter but provided excellent stlck-force-por-g characteristics for level flight stabilized high-rate tjrns. Stabilized turning performance Is presented In Figures 18 and 19 and compared to a 1-j level flight performance at the s.no conditions. The following grapn presents the maximum sustained bank angle, as limited by power available at mid C.G., and 7650 pounds and 9230 pounds gross weight and 4680 feet and 4750 feet density altitude respectively.

60 100 120 TRUE A;äSPE£P - KNOTS

FOR GrrlClAL USE ONLV

4.7 LEVEL ACCELERATIONS

Level flight accelerations were conducted at 7600 pounds and 9300 pounds at density altitudes of 3090 feet and 1750 feet respectively. Level accelerations were Initiated from stabilized flight at the speed near minimum power required, approximately 50 KCAS. Collective pitch was*set to obtain maximum power available. Altitude was maintained with cyclic pitch and the aircraft was allowed to accelerate to near maximum airspeed. Selection of maximum power was achieved by Increasing collective pitch as rapidly as possible to a value near maximum power. When engine speed and rotor speed transients were stabilized, an adjustment to maximum power was made. A coordinated nose-down attitude change accompanied the power Increase to maintain constant height. The aircraft was easy to control during acceleration from 50 KCAS to maximum airspeed. During the acceleration, collective pitch had to be continually Increased to maintain full power. Level flight acceleration data is presented In terms of energy per unit time (foot-pounds/second-dE/dt) versus true airspeed In Figures 21 and 22 and compared in Figure 20.

4.8 LEVEL DECELERATIONS

Level decelerations were accomplished at a test gross weight of approximately 9300 pounds from 112 KTAS to approximately 30 knots by reducing power to minimum practical and maintaining constant altitude. There were no limiting factors for this maneuver apart from the inherent drag characteristics and energy absorption capability of the rotor. As collective was reduced to minimum, the twist grip was closed to idle; this gave an 8- percent engine rotor needle split. Rotor speed was maintained between 95- percent and 100-percent RPM. Nose-up attitude was Increased to maintain height. At about 30 KCAS the twist grip was opened fully and power was increased to assist deceleration and establish a hover. The engine response was very good and twist grip control was precise and rapid. The aircraft displayed good flying qualities throughout this maneuver, although the nose- down trim change as collective was reduced was greater than allowed in Paragraphs 3.2.10.2 and 3.6.5.1, MIL-H-8501A (Reference e). Adequate control power, however, existed to compensate for this. The following plot of the true airspeed versus time presents the general deceleration performance at 9300 pounds and 1750 feet density altitude, at test day conditions.

10

a%: iise cuy DECELERATION PERFORMANCE. UOH-STAHPARD/ZED TEST DATA ROCATZr POPS W/TA/Ol/T COVE'S

GEAR UP 9300 LBS-DENSITY ALT.:I750PT

ZERO S/DESL/P GEAR U/* PEC EL EM A T/OA/

5/P£ FLARE: P£CELERAT/ON

TIME -SECONDS

As can be scon, a s l i g h t time advantage 1s gained using the side f l a r e uiJ iough t h i s i s not a recorrmended maneuver. This 1s discussed In paragraph 4 .9 .

4.9 LEVEL pr.CELSrtATIC:; SICE FLARES

Side f l a res were accomplished at a t es t gross weight of approximately i pOwiiui f r c i various entry airspeeds up to maximum obtainable airspeed.

7;.c tcchntquc useJ to executc t h i s naneuver was to maintain a s t r a i g h t path over the ground, decelerate and come to a hover w i th the a i r c r a f t heading perpendicular to o r i g i n a l f l i g h t path. In order to accomplish t h i s and remain w i t h i n the s i d e s l i p envelope of the UH-2B a i r c r a f t , the i n i t i a l bank was in the d i r e c t i o n o f f l i g h t u n t i l the f l i g h t path speed was d iss ipa ted enough to obtain a s i d e s l i p angle of 90 degrees. At t h i s po in t a sharp bank

11

FOa G'FiwAL USE ONLY

reversal was made and a hover was established without a change In altitude. This maneuver was highly Impractical because it was extremely difficult to remain within the operational sideslip envelope of the aircraft. Limiting factors for the UH-^B.or any other helicopter, were primarily those of the degree of difficulty In execution and the extremely uncoordinated nature of the maneuver. To maintain a straight path over the ground and to present maximum drag for deceleration, a severe sideslip had to be initiated simultaneously with a zero power deceleration (as in straight decelerations). Strong positive dihedral effect and directional stability at high speed necessitated large out-of-trlrn forces and caused extreme discomfort during the initial sideslip. The UH-2B had a large sideslip envelope and decelerated rapidly during this maneuver. The achievement of minimum stopping distances was limited only by pilot skill and persistence. Distance and time increased as a function of Increased airspeed, and In general an approximately 20-percent reduction In stoppinq time and distance could be obtained using this method as compared to time and distance required to stop during the straight deceleration. Side flares to toe right were less effective than to the left due to reduced rudder control power available needed to obtain high sideslip angles. Under conditions In which marginal power reserve for hovering was available the completion of the maneuver would have required more transient power than available. Under these conditions the maneuver would be impractical. The plot presented at test day conditions in Paragraph A.8 shows the reduction in time to perfonn a side flare deceleration as compared to a zero sideslip deceleration, with an entry true airspeed of 112 knots.

4„10 TURN REVERSALS

4.10.1 Turn reversals presented In Figure 23 were accomplished at 8800 pounds gross weight and at various airspeeds up to maximum airspeeds. The technique used was to stabilize on power, airspeed, and altitude, and rapidly displace the controls to obtain maximum useable roll rate. A maximum-effort level turn was held for 90 degrees and rapidly reversed for another 90 degrees. The maneuvering wa^ terminated wings level on original heading. Trim power and altitude were held constant throughout the maneuver.

4.10.2 The turn reversal maneuvers were entered at calibrated airspeeds of 101, 119, 127 and 128.5 knots. A smoothly coordinated roll and aft stick displacement were employed to initiate the turn. Ground references were used for heading control. Maximum load factor as limited by power and/or blade stall was used. Maximum comfortable roll rates of 47 degrees/second were obtained at 100 KCAS and reduced to 40 degrees/second at V . Blade stall during right rolls and during the turns was the main limixinq factor at V . Power available was the limiting factor at 101 and 119 KCAS. If more power were available during the turn at 101 KCAS and 119 KCAS, reduction in turning radius and in total time would be achieved. The following graph suimarizes the turn reversal performance at test dav conditions of 0800 pounds gross weight and 2850 feet density altitude.

12

i VUi üS£ uix

30

<<J> 70 1'

/20r

< '

ao -

2 \

6öL

00

60 L

so

< CO

< ^^

20

P5^ /5 ^H ^

/Ö SO 100 110 120 130

CALIBRATED AIRSPEED - KTS

As can bo soon from the presentation, the time required to complete the maneuver decreased with entry airspeed In the range of speeds tested.

4.11 TCAR DP.n? TUI^S •

4.11."i This niancuvcr consisted of possinrj over a pre-selectod icint on the qround, making a steep turn and returning over that spot In minimum time, inaintainimj a constant altitude. This maneuver was accomplished using various entry speeds up to maximum level flight speed. The entry calibrated airspeeds selected were 100, 120 and 133 knots. Optimum techniques for

13

-U C^lCiPi USE ONLY

obtaining the maximum rate and minimum radius turns at each entry speed were established prior to taking final test data. The following limitations were applicable during the maneuver:

Load Factor 3 g

Rotor Speed 96X - 100*

Bank Angle No limit

Torque No limit

Maximum Engine Speed 102%

Blade Stall 1/rev onset

4.11.2 As can be seen In the accompanying graph presented at test day conditions» as airspeed was decreased, time to complete the maneuver was also decreased within the speed range tested. Speeds at entry in excess of 100 KCAS were, therefore, reduced as quickly as possible to 100 knots during the initial roll after crossing the target. This was accomplished by reducing the collective pitch several Inches and Increasing load factor as the bank angle Increased. At 100 KCAS, full power was applied as Indicated by 102-percent engine speed and 1-percent rotor bleed; and maximum load factor was applied as limited by blade stall onset. Bank angle was adjusted to maintain constant height above the ground. As the target came back Into view, bank angle was reduced and maximum power was maintained to achieve the shortest flight path back to the target.

14

•J» »*. J

15

TEAR PROP TURNING PERFORMANCE

NOH-STANPAftP/2£ü T£ST DATA ROC/<£T POPS w/n/our CON£S

6R065 WEIGHT: 9500 L&S OtNölTY ALT 2400 FT • GtAR UP

100

90

80

70

60

50

KvT) k§

18

17

16

15

13

TIME TO TU£ LETT AVERAGED 2 SEC. LESS TUAAI TO THE RIGUT

80 100 110 120 130 140 ENTRY CALIBRATED AIRSPEED-KTS

15

4.11.3 Peak roll rates of approximately 40 degrees/second and maximum bank anales between 70 degrees and 80 degrees were used on entry. Load factor (g) was held between 2.3 and 2.7 g. The time to return to target depended to a great extent upon consistency of pilot technique but. In general, varied from 15 seconds at 100 KCAS entry speed to 17 seconds at 133J(CAS entry speed. Blade stall on the UH-2B was the main limiting factor*1n this type of maneuver, particularly at high gross weight. During the maximum g portion of the turn, there was very little warning of the onset of blade stall and. In spite of attempts to avoid It, blade stall was reached on almost every occasion at 9300 pounds. Blade stall on the UH-2B was characterized by heavy 1-per-rev vibration, longitudinal pitch-up, and Increased 4-per-rev vibration. Once encountered, blade stall was self sustaining and required positive correction by reducing power and decreasing load factor (applying forward stick). A positive aft stick force of approximately 10 pounds was required to hold 2.8 g and a relaxation of this force aided In the recovery from blade stall. It Is stressed that the ease of reaching blade stall and lack of stall warning are very undesirable featuresln the UH-2B helicopter. The maximum-rate, mlmmum-raoius turn maneuver, however, was not the worst case with respect to maneuvering stability. The stability characteristics during diving pull-outs are discussed In Paragraph 4.13.4, MANEUVERING STABILITY. Blade stall was also Induced by large roll demands to the right during the entry and exit rolls of the tear drop. A reduction In lateral cyclic control Input was required to recover from the blade stall.

4.12 AUTORDTATIONAL ENTRIES

4.12.1 Level flight autorotatlonal entries were conducted at a test gross weight of 9300 pounds at various airspeeds up to maximum airspeed by simulating total power failure. A 1-second time delay was used during this maneuver; I.e., the procedure was to close the throttle, hold all controls fixed for 1 second, then Initiate recovery to attain best autorotatlve airspeed In minimum time.

4.12.2 The technique for the UH-2B was critical with regard to maximum height gain after a 1-second delay. If the collective was lowered fully at the same time the flare was Initiated, I.e., after 1 second, only 75 to 100 feet were gained at an entry speed of 130 KCAS. If collective was held fixed or reduced only slightly after the 1-second delay and a rapid flare was made, nearly 200 feet of altitude were gained.

4.12.3 Rotor decay following power loss dictated the length of time that the collective could be held up during the flare and, therefore, the height gained. When rotor RPM had decayed to 90 percent, collective was reduced to maintain at least this value. At the peak of the height gain, the aircraft speed was nearly down to best autorotatlonal speed. The following graph is a sunwary of autorotatlonal entry test day data for the UH-2B at approximately 9300 pounds gross weight and 2400 feet density altitude.

16

FOR WrWl m ONLY

AUTOROTATION AL ENTRY NON-STANPAßPJZ£D T£ST DATA GEAR UP UNTIL G5 kT6 REACME.D ROCKET P0D5 WITWOUT CONES

I SECOND DELAY GROSS WT'. 9500 L&5 / DENSITY ALT••2400 FT.

p

18

16

14

I?.

10

8

I

AVERAGE TIME f ROM TUROTTLE CUOP TO RETURH TOWfWALTy

T c

TIME TO 65 KT^S ^ PEAK HEJGUT

^1

ZOO

100 -

0 100 110 120 130 140

ENTRV TRUE AIRSPEED-KNOTS

17

« * I » 4 i fl

4.13 STABILITY AND CONTROL

4.13J STATIC LONGITUDINAL SPEED STABILITY

Collective-fixed static longitudinal speed stability tests were conducted at calibrated airspeeds of 61, 100 and 131 knots at 9000-pounds gross weight and density altitude of 3500 feet. Test data presented In Figure 31 shows the static longitudinal stability about trim versus calibrated airspeed to be slightly positive at 60 and 100 KCAS, becoming neutral at 130 KCAS. The control position trim curves In Figures 32 and 33 show that the trimmed longitudinal control position was positive up to nearly maximum calibrated airspeed, then became neutral for 277 rotor RPM and negative for 268 rotor RPM. This condition existed at conditions tested between 7600 and 9690 pounds gross weight, 4000 feet to 6350 feet density altitude at an approximately mid CG. The "bob weight" In the longitudinal control system was noticeable In rough air since It applied a forward or aft stick force as a function of positive or negative normal acceleration. If the pilot allowed the stick to move, the "bob weight" corrected for gusts quite well. A large longitudinal trim change resulted from changes In collective position which, although In excess of requirements of paragraph 3.5.5.1, MIL-H-8501A (Reference e), were partly compensated for by the "bob weight."

4.13.2 STATIC LATERAL DIRECTIONAL STABILITY

Static lateral directional stability tests were conducted at calibrated airspeeds of 50, 100 and 127 knots and density altitudes of 3800 and 3900 feet. A rotor speed of 277 RPM was used at an approximately mid CG. Test results, presented In Figures 34 through 36, show that the static lateral directional stability was strongly positive but that a nonsymmetrlcal longitudinal trim change accompanied the sideslip. Under transient conditions the "bob weight" assisted In compensating for this characteristic.

4.13.3 CONTROLLABILITY

4.13.3.1 Roll response characteristics, presented In Figures 37 through 39 with a time history of a left lateral step In Figure 40, were investigated at calibrated airspeeds of 50, 100 and 128 knots at approximately 9300 pounds gross weight and an approximately mid CG. Lateral control response was characterized by high roll rate damping. In most cases thß roll rate was over-damped. At maximum airspeeds roll rate response was approximately 10 degrees/second/inch or 5 degrees/second/pound of force, which was lesi than optimum. Responses after 1 second for 1-Inch and 2-Inch inputs were approximately 8 degrees and 15 degrees respectively. For the armed mission, roll response of 15 to 20 degrees/second/inch with maximum roll rates of approximately 50 degrees to 70 degrees/second Is desirable.

4.13.3.2 Qualitatively, the longitudinal control response was satisfactory. The response was similar in sensitivity, damping and rate to that of most

18

helicopters of this class. During maneuvering flight the effect of the "bob weight" was very noticeable In the longitudinal axis, where it caused a stick force proportional to g. This force, although unusual when first experienced. Improved maneuvering stability.

4.13.3.3 Qualitatively, the directional control response was heavily damped and pedal forces were high. Due to the longitudinal trim change associated with sideslip, directional control step Inputs caused a nose-down pitch with right pedal and a small nose-up pitch with left pedal. This characteristic Is undesirable for weapons firing.

4.13.4 MANEUVERING STABILITY

4.13.4.1 Maneuvering stability characteristics were Investigated In terms of stick position and stick force per g. Standard flight test techniques were used. Plots of stick position (stick-fixed) and stick force (stick- free) per g are presented In Figures 41 and 42. The aircraft was evaluated with the ASE OFF at average gross weights of 7900 pounds and 9130 pounds at density altitudes of 3000 feet and 4000 feet respectively. The primary technique used was the synmetrlcal pull-up since this was the most critical case. Wind-up turns and spiral dives were qualitatively checked and In both cases larger maneuver margins were apparent. The ASE was also qualitatively checked and Increased maneuver margins were apparent with the ASE ON.

4.13.4.2 The maneuvering stability test results at 7900 pounds gross weight. In terms of stick force versus g a^d stick displacement from trim versus g are presented In Figure 41. The results in both cases show a decreasing maneuver margin with increasing airspeed. Positive stick-fixed maneuver margins are shown at 80 KCAS and 100 KCAS, but at 120 KCAS and 140 KCAS neutral and negative stick-fixed margins respectively are shown. This indicates that as g was apolied the rotor disc was flapping back, and that, after Initiating the pull-up with aft stick, forward stick had to be applied to prevent q from Increasing above the desired value. Thus the "bob weight" effectiveness was reduced and the resulting stick force per g was reduced. Results of the stick-free maneuver margin tests are similar to the stick- fixed results in that at 80 and 100 KCAS a positive stick force per g exists which then becomes neutral at 120 KCAS and negative at 140 KCAS.

4.13.403 The effect of increased weight on maneuvering stability was determined by repeating the tests at 9130 pounds. These test results are shown In Figure 42. The same trend as in the lightweight tests Is apparent, but the speeds at which significant reductions in maneuvering stability occur are lower. Stick force per g as a function of g also decreases at a greater rate and at lower g values. This would be explained by the earlier onset of blade stall at the heavier weights. At 120 KCAS and 9130 pounds (See Figure 42), negative stick-fixed stability exists at all values of g. Thus, If the stick is held In the aft displaced position, a divergent pitch-up

19

results (See Figure 43, time history of an aft longitudinal step). This continues until severe blade stall is reached or recovery action is initiated. As shown in Figure 43, the divergence continues in spite of slow forward stick motion. Note that aft stick force Increases slightly under the Influence of the "bob weight" and that, when the stick Is released, corrective forward stick Is applied. Collective pitch reduction is necessary, however, due to the severity of the blade stall. This characteristic is prohibit id by Paragraph 3.2.8, MIL-H-8501A (Reference e). Inadequate warning of the pitch- up exists. This Indicates that a limited dive recovery envelope should be Imposed,

4.13.4.4 Qualitatively, under conditions where speeds below 115 KCAS and weights below 8500 pounds were used, the maneuvering stability of the UH-2B was good up to 2.5 g. The "boo weight" provided excellent stick-force-per-g characteristics and high load factors could be used with confidence and precision. The trend of stick-free maneuvering stability with increased g, weight, and airspeed was one of reduced maneuvering stability. The problem appeared to be that of Increasing unstable rotor flap-back as blade stall was approached, causing negative stick position per g. The "bob weight" force was, therefore, canceled out by the pilot's having to allow the stick to move forward. This effect became very pronounced during symmetrical diving pull-outs at heavy gross weiahts and high speeds that resulted In a divergent longitudinal pitch-up in which load factor rapidly built up until heavy blade stall was encountered and immediate recovery was essential. In turning flight the effect was less pronounced.

4.13.5 DYNAMIC LATERAL DIRECTIONAL STABILITY

Limited dynamic lateral directional stability tests were conducted at 120 KCAS, 9280 pounds and 2900 feet density altitude. Pedal releases were made from a 10-degree sideslip and time histories are presented in Figures 44 and 45. As can be seen in Figure 44, aircraft response with the yaw damper ON from a 10-degree release is "dead beat." Figure 45 shows that with a yaw damper OFF, approximately K7 seconds and 1.5 cycles are required for the aircraft to damp to half amplitude. The use of the vaw damper Increases the effect!veriiass of tne aircraft as a weapons platform.

SECTION 2. APPENDICES

APPENDIX I

TEST DATA

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•APPLNDIX II

. REFERENCES

a. (C) Message 13923. AMCRD, Hq, U. S. Army Materiel Command (USAMC), 23 October 1965, subject: Expedited Flight Test Evaluation (U).

b. (C) Message 14297, AMCRD, Hq, USAMC, 28 October 1965, subject: Addition of UH-2 to Expedited Flight Test Evaluation (U).

c. (C) Plan of Test of Armed Helicopters (U), U. S. Army Aviation Test Activity, 28 October 1965.

d. (C) Report, "Early Engineering Confirmatory Flight Test Evaluation Improved Armed Helicopter," (U), Hq, USAMC, undated.

e. Military Specification MII-H-8501A, "General Requirements for Helicopter Flying and Ground Handling Qualities," 7 September 1961.

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